Gas turbine engine shaft bearing configuration

ABSTRACT

A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft.

This application is a continuation of U.S. application Ser. No.13/362,170, filed on Jan. 31, 2012.

BACKGROUND

This disclosure relates to a gas turbine engine bearing configurationfor a shaft. In one example, the bearing arrangement relates to a lowshaft.

A typical jet engine has two or three spools, or shafts, that transmittorque between the turbine and compressor sections of the engine. Eachof these spools is typically supported by two bearings. One bearing, forexample, a ball bearing, is arranged at a forward end of the spool andis configured to react to both axial and radial loads. Another bearing,for example, a roller bearing is arranged at the aft end of the spooland is configured to react only to radial loads. This bearingarrangement fully constrains the shaft except for rotation, and axialmovement of one free end is permitted to accommodate engine axialgrowth.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a corehousing that has an inlet case and an intermediate case thatrespectively provide an inlet case flow path and an intermediate caseflow path. The shaft supports a compressor section that is arrangedaxially between the inlet case flow path and the intermediate case flowpath. The shaft includes a main shaft and a flex shaft having bellows.The flex shaft is secured to the main shaft at a first end and includesa second end opposite the first end. A geared architecture is coupled tothe shaft, and a fan coupled to and rotationally driven by the gearedarchitecture. The geared architecture includes a sun gear supported onthe second end. A first bearing supports the shaft relative to theintermediate case and a second bearing supporting the shaft relative tothe inlet case. The second bearing is arranged radially outward from theflex shaft.

In a further embodiment of any of the above, the shaft includes a hubsecured to the main shaft. The compressor section includes a rotormounted to the hub. The hub supports the second bearing.

In a further embodiment of any of the above, the inlet case includes aninlet case portion defining the inlet case flow path. A bearing supportportion is removably secured to the inlet case portion. The secondbearing is mounted to the bearing support portion.

In a further embodiment of any of the above, the inlet case includes afirst inlet case portion defining the inlet case flow path. A bearingsupport portion is removably secured to the inlet case portion. Thesecond bearing is mounted to the bearing support portion.

In a further embodiment of any of the above, the intermediate caseincludes an intermediate case portion defining the intermediate caseflow path. A bearing support portion is removably secured to theintermediate case portion. The first bearing is mounted to the bearingsupport portion.

In a further embodiment of any of the above, the first bearing is a ballbearing and the second bearing is a roller bearing.

In a further embodiment of any of the above, the first and secondbearings are arranged in separate sealed lubrication compartments.

In a further embodiment of any of the above, there is a lubricationcompartment. The second bearing and the geared architecture are arrangedin the lubrication compartment.

In one exemplary embodiment, a gas turbine engine includes a corehousing. The core housing includes an inlet case and an intermediatecase that respectively provide an inlet case flow path and anintermediate case flow path. A shaft supports a compressor section thatis arranged axially between the inlet case flow path and theintermediate case flow path. The compressor section includes a variablestator vane array. A first bearing supports the shaft relative to theintermediate case and a second bearing supports the shaft relative tothe inlet case. The second bearing is axially aligned with and radiallyinward of the variable stator vane array. A geared architecture iscoupled to the shaft. A fan is coupled to and rotationally driven by thegeared architecture. The shaft includes a main shaft and a flex shafthaving bellows. The flex shaft is secured to the main shaft at a firstend and includes a second end opposite the first end. The gearedarchitecture includes a sun gear supported on the second end, and thesecond bearing is arranged radially outward from the flex shaft. Theshaft includes a hub secured to the main shaft. The compressor sectionincludes a rotor mounted to the hub. The hub supports the secondbearing.

In one exemplary embodiment, a gas turbine engine includes a corehousing that provides a core flow path. The gas turbine engine includesa fan and a shaft that supports a compressor section arranged within thecore flow path. The compressor section is fluidly connected to the fan.The compressor section includes a first pressure compressor and a secondpressure compressor upstream from the first pressure compressor. Thesecond pressure compressor includes multiple compressor stages. Firstand second bearings support the shaft relative to the core housing andare arranged radially inward of and axially overlapping with at leastsome of the multiple compressor stages. The gas turbine engine is a highbypass geared aircraft engine having a bypass ratio of greater thanabout six (6).

In a further embodiment of any of the above, combustor is fluidlyconnected to the compressor section. A turbine section is fluidlyconnected to the combustor. The turbine section includes a high pressureturbine and a low pressure turbine.

In a further embodiment of any of the above, the core housing includes afirst inlet case portion defining an inlet case flow path, and a bearingsupport portion removably secured to the inlet case portion. A secondbearing mounts to the bearing support portion.

In a further embodiment of any of the above, the core housing includesan intermediate case portion defining an intermediate case flow path,and a bearing support portion removably secured to the intermediate caseportion. The first bearing is mounted to the bearing support portion.

In a further embodiment of any of the above, the multiple compressorstages include a variable stator vane array, rotatable compressorblades, and a fixed stator vane array.

In one exemplary embodiment, a gas turbine engine includes a corehousing that provides a core flow path. The gas turbine engine alsoincludes a fan and a shaft that supports a compressor section arrangedwithin the core flow path. The compressor section is fluidly connectedto the fan. The compressor section includes a first pressure compressorand a second pressure compressor upstream from the first pressurecompressor. The second pressure compressor includes multiple compressorstages. The first and second bearings support the shaft and are relativeto the core housing and are arranged radially inward of and axiallyoverlapping with at least some of the multiple compressor stages. Acombustor is fluidly connected to the compressor section. A turbinesection is fluidly connected to the combustor. The turbine sectionincludes a high pressure turbine and a low pressure turbine. The gasturbine engine includes at least one of a low Fan Pressure Ratio of lessthan about 1.45 and a low pressure turbine pressure ratio that isgreater than about 5.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine.

FIG. 2 is a cross-sectional view of a front architecture of the gasturbine engine shown in FIG. 1.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath B whilethe compressor section 24 drives air along a core flowpath C (as shownin FIG. 2) for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57supports one or more bearing systems 38 in the turbine section 28. Theinner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A, whichis collinear with their longitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example a high-bypass geared aircraft engine. In afurther example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a star gear systemor other gear system, with a gear reduction ratio of greater than about2.3 and the low pressure turbine 46 has a pressure ratio that is greaterthan about 5. In one disclosed embodiment, the engine 20 bypass ratio isgreater than about ten (10:1), the fan diameter is significantly largerthan that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to inletof low pressure turbine 46 as related to the pressure at the outlet ofthe low pressure turbine 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a star gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.5:1. It should be understood, however, that the above parametersare only exemplary of one embodiment of a geared architecture engine andthat the present invention is applicable to other gas turbine enginesincluding direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as bucket cruiseThrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)] ^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

Referring to FIG. 2, a core housing 60 includes an inlet case 62 and anintermediate case 64 that respectively provide an inlet case flowpath 63and a compressor case flowpath 65. Together, the inlet and compressorcase flowpaths 63, 65, in part, define a core flowpath through theengine 20, which directs a core flow C.

The intermediate case 64 includes multiple components, which includesthe intermediate case portions 66, and the bearing support 68 in theexample, which are removably secured to one another. The bearing supportportion 68 has a first bearing 70 mounted thereto, which supports theinner shaft 40 for rotation relative to the intermediate case 64. In oneexample, the first bearing 70 is a ball bearing that constrains theinner shaft 40 against axial and radial movement at a forward portion ofthe inner shaft 40. The first bearing 70 is arranged within a bearingcompartment 71.

In the example, the inner shaft 40 is constructed of multiple componentsthat include, for example, a main shaft 72, a hub 74 and a flex shaft76, which are clamped together by a nut 80 in the example. The firstbearing 70 is mounted on the hub 74. The flex shaft 76 includes firstand second opposing ends 82, 84. The first end 82 is splined to the hub74, and the second end 84 is splined to and supports a sun gear 86 ofthe geared architecture 48. Bellows 78 in the flex shaft 76 accommodatevibration in the geared architecture 48.

The geared architecture includes star gears 88 arrangedcircumferentially about and intermeshing with the sun gear 86. A ringgear 90 is arranged circumferentially about and intermeshes with thestar gears 88. A fan shaft 92 is connected to the ring gear 90 and thefan 42 (FIG. 1). A torque frame 94 supports the star gears 88 andgrounds the star gears 88 to the housing 60. In operation, the innershaft 40 rotationally drives the fan shaft 92 with the rotating ringgear 90 through the grounded star gears 88.

The low pressure compressor 44 includes multiple compressor stagesarranged between the inlet and intermediate case flowpaths 63, 65, forexample, first and second compressor stages 98, 100, that are secured tothe hub 74 by a rotor 96. The first bearing 70 is axially aligned withone of the first and second compressor stages 98, 100. In one example, avariable stator vane array 102 is arranged upstream from the first andsecond compressor stages 98, 100. Struts 104 are arranged upstream fromthe variable stator vane array 102. An array of fixed stator vanes 106may be provided axially between the first and second compressor stages98, 100. Although a particular configuration of low pressure compressor44 is illustrated, it should be understood that other configurations maybe used and still fall within the scope of this disclosure.

The inlet case 62 includes inlet case portions 108, and bearing support110, which are removably secured to one another. The bearing supportportion 110 and torque frame 94 are secured to the inlet case portion108 at a joint 109. The bearing support portion 110 supports a secondbearing 112, which is a rolling bearing in one example. The secondbearing 112 is retained on the hub 74 by a nut 113, for example, and isarranged radially outward from the flex shaft 76 and radially betweenthe torque frame 94 and flex shaft 76. In the example, the secondbearing 112 is axially aligned with and radially inward of the variablestator vane array 102. The geared architecture 48 and the second bearing112 are arranged in a lubrication compartment 114, which is separatefrom the bearing compartment 71 in the example.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: a core housingincluding an inlet case and an intermediate case that respectivelyprovide an inlet case flow path and an intermediate case flow path; ashaft supporting a compressor section that is arranged axially betweenthe inlet case flow path and the intermediate case flow path, whereinthe shaft includes a main shaft and a flex shaft having bellows, theflex shaft secured to the main shaft at a first end and including asecond end opposite the first end; a geared architecture coupled to theshaft, and a fan coupled to and rotationally driven by the gearedarchitecture, wherein the geared architecture includes a sun gearsupported on the second end; a first bearing supporting the shaftrelative to the intermediate case; and a second bearing supporting theshaft relative to the inlet case, and the second bearing is arrangedradially outward from the flex shaft.
 2. The gas turbine engineaccording to claim 1, wherein the shaft includes a hub secured to themain shaft, and the compressor section includes a rotor mounted to thehub, the hub supporting the second bearing.
 3. The gas turbine engineaccording to claim 2, wherein the inlet case includes an inlet caseportion defining the inlet case flow path, and a bearing support portionremovably secured to the inlet case portion, the second bearing mountedto the bearing support portion.
 4. The gas turbine engine according toclaim 1, wherein the inlet case includes a first inlet case portiondefining the inlet case flow path, and a bearing support portionremovably secured to the inlet case portion, the second bearing mountedto the bearing support portion.
 5. The gas turbine engine according toclaim 1, wherein the intermediate case includes an intermediate caseportion defining the intermediate case flow path, and a bearing supportportion removably secured to the intermediate case portion, the firstbearing mounted to the bearing support portion.
 6. The gas turbineengine according to claim 1, wherein the first bearing is a ball bearingand the second bearing is a roller bearing.
 7. The gas turbine engineaccording to claim 1, wherein the first and second bearings are arrangedin separate sealed lubrication compartments.
 8. The gas turbine engineaccording to claim 1, comprising a lubrication compartment, the secondbearing and the geared architecture arranged in the lubricationcompartment.
 9. A gas turbine engine comprising: a core housingincluding an inlet case and an intermediate case that respectivelyprovide an inlet case flow path and an intermediate case flow path; ashaft supporting a compressor section that is arranged axially betweenthe inlet case flow path and the intermediate case flow path, whereinthe compressor section includes a variable stator vane array; a firstbearing supporting the shaft relative to the intermediate case; a secondbearing supporting the shaft relative to the inlet case, and the secondbearing is axially aligned with and radially inward of the variablestator vane array; and a geared architecture coupled to the shaft, and afan coupled to and rotationally driven by the geared architecture,wherein the shaft includes a main shaft and a flex shaft having bellows,the flex shaft secured to the main shaft at a first end and including asecond end opposite the first end, wherein the geared architectureincludes a sun gear supported on the second end, and the second bearingis arranged radially outward from the flex shaft, wherein the shaftincludes a hub secured to the main shaft, and the compressor sectionincludes a rotor mounted to the hub, the hub supporting the secondbearing.
 10. A gas turbine engine comprising: a core housing providing acore flow path; a fan; a shaft supporting a compressor section arrangedwithin the core flow path, wherein the compressor section is fluidlyconnected to the fan, the compressor section comprising a first pressurecompressor and a second pressure compressor upstream from the firstpressure compressor, the second pressure compressor including multiplecompressor stages ; and first and second bearings supporting the shaftrelative to the core housing and are arranged radially inward of andaxially overlapping with at least some of the multiple compressorstages; and wherein the gas turbine engine is a high bypass gearedaircraft engine having a bypass ratio of greater than about six (6). 11.The gas turbine engine according to claim 10, further comprising: acombustor fluidly connected to the compressor section; a turbine sectionfluidly connected to the combustor, the turbine section comprising: ahigh pressure turbine; and a low pressure turbine.
 12. The gas turbineengine according to claim 10, wherein the core housing includes a firstinlet case portion defining an inlet case flow path, and a bearingsupport portion removably secured to the inlet case portion, the secondbearing mounted to the bearing support portion.
 13. The gas turbineengine according to claim 10, wherein the core housing includes anintermediate case portion defining an intermediate case flow path, and abearing support portion removably secured to the intermediate caseportion, the first bearing mounted to the bearing support portion. 14.The gas turbine engine according to claim 10, wherein the multiplecompressor stages includes a variable stator vane array, rotatablecompressor blades, and a fixed stator vane array.
 15. A gas turbineengine comprising: a core housing providing a core flow path; a fan; ashaft supporting a compressor section arranged within the core flowpath, wherein the compressor section is fluidly connected to the fan,the compressor section comprising a first pressure compressor and asecond pressure compressor upstream from the first pressure compressor,the second pressure compressor including multiple compressor stages; andfirst and second bearings supporting the shaft relative to the corehousing and being arranged radially inward of and axially overlappingwith at least some of the multiple compressor stages a combustor fluidlyconnected to the compressor section; a turbine section fluidly connectedto the combustor, the turbine section comprising: a high pressureturbine; a low pressure turbine; and wherein the gas turbine engineincludes at least one of a low Fan Pressure Ratio of less than about1.45 and a low pressure turbine pressure ratio that is greater thanabout 5.